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At least 523 records · Page 29

All sources NASA NTRS DOE OSTI

NASA NTRS2016

Additively Manufactured Combustion Devices Components for LOX/Methane Applications

Marshall Space Flight Center (MSFC) has designed, fabricated, and hot-fire tested a variety of successful injectors, chambers, and igniters for potential liquid oxygen (LOX) and methane (CH4) systems since 2005. The most recent efforts have focused on components with additive manufacturing (AM) to include unique design features, minimize joints, and reduce final machining efforts. Inconel and copper alloys have been used with AM processes to produce a swirl coaxial injector and multiple methane cooled thrust chambers. The initial chambers included unique thermocouple ports for measuring local coolant channel temperatures along the length of the chamber. Results from hot-fire testing were used to anchor thermal models and generate a regeneratively cooled thruster for a 4,000 lbf LOX/CH4 engine. The completed thruster will be hot-fire tested in the summer of 2016 at MSFC. The thruster design can also be easily scaled and used on a 25,000 lbf engine. To further support the larger engine design, an AM gas generator injector has been designed. Hot-fire testing on this injector is planned for the summer of 2016 at MSFC.

Greene, Sandra Elam↗

DOE OSTI2024

Warm Unit Stand Assembly and Testing for Proton Improvement Plan - II

The Proton Improvement Project II (PIP II), situated at Fermilab, signifies a critical advancement in the realm of high-energy physics. PIP II, a linear accelerator renovation endeavor, aspires to double the output of the existing proton beam, reinforcing Fermilab’ s position as a vanguard in this field. The proton beam's creation involves a series of high-frequency cryomodules, necessitating environmental temperature arrays of magnets and vacuum pumps henceforth described as ‘Warm Units.’ The Warm Units include design features of : 1. insertion between the cryomodules, 2. multi-stage adjustments in six degrees of freedom to ensure precise beam alignment; 3. Natural frequency modes that avoid specified ranges. In adherence to Fermilab’ s stringent safety standards, comprehensive testing procedures are slated to validate the Warm Units' capacity to simultaneously achieve local alignment adjustments while averting low-frequency oscillation less than 15 Hertz. The testing regimen will scrutinize the Warm Units' alignment adjustment capabilities and their resilience against low-frequency oscillations, with an overarching objective of satisfying the Final Design Review criteria specific to the PIP II initiative. This project works to furnish technicians responsible for the assembly of Warm Units with a comprehensive guide for successful installation, concurrently documenting any requisite design modifications essential for achieving a real-world configuration.

43 PARTICLE ACCELERATORS↗

NASA NTRS1980

Model aerodynamic test results for two variable cycle engine coannular exhaust systems at simulated takeoff and cruise conditions

Wind tunnel tests were conducted to evaluate the aerodynamic performance of a coannular exhaust nozzle for a proposed variable stream control supersonic propulsion system. Tests were conducted with two simulated configurations differing primarily in the fan duct flowpaths: a short flap mechanism for fan stream control with an isentropic contoured flow splitter, and an iris fan nozzle with a conical flow splitter. Both designs feature a translating primary plug and an auxiliary inlet ejector. Tests were conducted at takeoff and simulated cruise conditions. Data were acquired at Mach numbers of 0, 0.36, 0.9, and 2.0 for a wide range of nozzle operating conditions. At simulated supersonic cruise, both configurations demonstrated good performance, comparable to levels assumed in earlier advanced supersonic propulsion studies. However, at subsonic cruise, both configurations exhibited performance that was 6 to 7.5 percent less than the study assumptions. At take off conditions, the iris configuration performance approached the assumed levels, while the short flap design was 4 to 6 percent less.

Nelson, D. P.↗

NASA NTRS1976

Television applications of interline-transfer CCD arrays

The design features and characteristics of interline transfer (ILT) CCD arrays with 190 x 244 and 380 x 488 image elements are reviewed, with emphasis on optional operating modes and system application considerations. It was shown that the observed horizontal resolution for a TV system using an ILT image sensor can approach the aperture response limit determined by photosensor site width, resulting in enhanced resolution for moving images. Preferred camera configurations and read out clocking modes for maximum resolution and low light sensitivity are discussed, including a very low light level intensifier CCD concept. Several camera designs utilizing ILT-CCD arrays are described. These cameras demonstrate feasibility in applications where small size, low-power/low-voltage operation, high sensitivity and extreme ruggedness are either desired or mandatory system requirements.

Hoagland, K. A.↗

NASA NTRS1993

Heat pipe cooled heat rejection subsystem modelling for nuclear electric propulsion

NASA LeRC is currently developing a FORTRAN based computer model of a complete nuclear electric propulsion (NEP) vehicle that can be used for piloted and cargo missions to the Moon or Mars. Proposed designs feature either a Brayton or a K-Rankine power conversion cycle to drive a turbine coupled with rotary alternators. Both ion and magnetoplasmodynamic (MPD) thrusters will be considered in the model. In support of the NEP model, Rocketdyne is developing power conversion, heat rejection, and power management and distribution (PMAD) subroutines. The subroutines will be incorporated into the NEP vehicle model which will be written by NASA LeRC. The purpose is to document the heat pipe cooled heat rejection subsystem model and its supporting subroutines. The heat pipe cooled heat rejection subsystem model is designed to provide estimate of the mass and performance of the equipment used to reject heat from Brayton and Rankine cycle power conversion systems. The subroutine models the ductwork and heat pipe cooled manifold for a gas cooled Brayton; the heat sink heat exchanger, liquid loop piping, expansion compensator, pump and manifold for a liquid loop cooled Brayton; and a shear flow condenser for a K-Rankine system. In each case, the final heat rejection is made by way of a heat pipe radiator. The radiator is sized to reject the amount of heat necessary.

Moriarty, Michael P.↗

NASA NTRS1992

Design, analysis, and testing of the Phase 1 CSI Evolutionary Model erectable truss

This report addressed the design, analysis, and testing of the erectable truss structure for the Phase 1 CSI Evolutionary Model (CEM) testbed. The Phase 1 CEM testbed is the second testbed to form part of an ongoing program of focused research at NASA/LaRC in the development of Controls-Structures Integration (CSI) technology. The Phase 1 CEM contains the same overall geometry, weight, and sensor locations as the Phase 0 CEM, but is based in an integrated controller and structure design, whereby both structure and controller design variables are sized simultaneously. The Phase 1 CEM design features seven truss sections composed of struts with tailored mass and stiffness properties. A common erectable joint is used and the strut stiffness is tailored by varying the cross-sectional area. To characterize the structure, static tests were conducted on individual struts and 10-bay truss assemblies. Dynamic tests were conducted on 10-bay truss assemblies as well as the fully-assembled CEM truss. The results indicate that the static and dynamic properties of the structure are predictable, well-characterized, and within the performance requirements established during the Phase 1 CEM integrated controller/structure design analysis.

Gronet, M. J.↗

NASA NTRS2001

A Low Temperature, Reverse Brayton Cryocooler

This status report covers the fifty-second month of a project to develop a low temperature, reverse-Brayton cryocooler using turbomachines. This program consists of a Basic Phase and four Option Phases. Each of the Phases is directed to a particular load/temperature combination. The technology and fundamental design features of the components used in these systems are related but differ somewhat in size, speed, and some details in physical geometry. Each of the Phases can be carried out independently of the others, except that all of the Phases rely on the technology developed and demonstrated during the Basic Phase. The Basic Phase includes the demonstration of a critical component and the production of a prototype model cryocooler. The critical technology demonstration will be the test of a small turboalternator over a range of conditions at temperatures down to 6 K. These tests will provide design verification data useful for the further design of the other coolers. The prototype model cooler will be designed to provide at least 5 mW of cooling at 6 K. The heat rejection temperature for this requirement is 220 K or greater. The input power to the system at these conditions is to be less than 60 W.

Swift, Walter L.↗

NASA NTRS2010

The Landsat Data Continuity Mission Operational Land Imager (OLI) Radiometric Calibration

The Operational Land Imager (OLI) on the Landsat Data Continuity Mission (LDCM) has a comprehensive radiometric characterization and calibration program beginning with the instrument design, and extending through integration and test, on-orbit operations and science data processing. Key instrument design features for radiometric calibration include dual solar diffusers and multi-lamped on-board calibrators. The radiometric calibration transfer procedure from NIST standards has multiple checks on the radiometric scale throughout the process and uses a heliostat as part of the transfer to orbit of the radiometric calibration. On-orbit lunar imaging will be used to track the instruments stability and side slither maneuvers will be used in addition to the solar diffuser to flat field across the thousands of detectors per band. A Calibration Validation Team is continuously involved in the process from design to operations. This team uses an Image Assessment System (IAS), part of the ground system to characterize and calibrate the on-orbit data.

Markham, Brian L.↗

DOE OSTI2023

Radiation Damage Analysis of FNSF Components Using McCad and MCNP

The Fusion Energy System Studies Fusion Nuclear Science Facility (FESS-FNSF) concept represents a transitional step between ITER and a commercial fusion power plant. The FNSF is a conceptualized D-T fueled tokamak with 518 MW of fusion power that has been extensively used to explore and optimize design features. The energetic 14.1-MeV neutrons can produce significant localized heating and activations, and can cause damage to plasma-facing components, which can determine maintenance/outage scheduling needs and also impact the lifetime of the device as a whole. This study illustrates a neutronics analysis that was conducted on a 22.5-degree symmetric sector of the FNSF with the goal of understanding the neutron heating and radiation damage that can be characterized by quantifying the displacements per atom (dpa). Concurrently, this study also focused on the development of analysis capabilities by converting a three-dimensional computer-aided design model of the FNSF into MCNP6.2 input using the McCad code. Accordingly, some confirmatory results on tritium production and the tritium breeding ratio (TBR) are provided to support model validation. The results produced by MCNP6.2 simulations showed that the highest heating and damage occurred in the outboard region, which concentrated approximately 290 MW of the total nuclear heating, in contrast to 97 MW within the inboard region. These results are consistent with previous studies that employed earlier versions of the FNSF concept and different modeling approaches. Furthermore, this study also provides additional details on neutron wall loading, as well as total heating from neutrons and gammas, results which show the total heating of the device (16 sectors) is approximately 477.83 ± 0.80% MW, indicating a neutron energy multiplication factor of 1.15. Additionally, the capability to calculate hydrogen and helium production, as well as dpa, is illustrated. Finally, the neutronics effects of using alternative materials to tungsten carbide were evaluated for the vacuum vessel, low-temperature shield, and structural ring components, which showed that compounds like YH 2 , Mg(BH 4 ) 2 , and ZrH 2 could reduce the total heating on the magnet and also reduce the TBR.

70 PLASMA PHYSICS AND FUSION TECHNOLOGY↗

DOE OSTI2025

A Combined Computational and Mathematical Analysis of Interconnect Fatigue Potential in Photovoltaic Modules

A finite element model of a 60-cell monocrystalline silicon glass-polymer photovoltaic module was simulated with ±1.0 kPa and ±2.4 kPa loads applied to the glass to calculate the deformation under load. Cell-to-cell displacements were used to approximate interconnect strain and stress. A mathematical fatigue cycle life relation was fitted to data for the interconnect material (copper), to generate a life prediction at each interconnect location based on the local stress means, reversal extents, and amplitudes. Interconnect stress was found to be significantly asymmetric about zero despite symmetric positive and negative module loads due to laminate thickness offsets about the neutral plane and the effects of module framing. Cycle life results indicated that interconnect fatigue failure was unlikely to occur over a 30-year lifetime of conservative wind and snow load cycles since the typical cell design feature of leaving some unconstrained length between the cell edge and first solder pad increases the effective gauge length and decreases the stress levels below the material endurance limit. Follow-up analyses found that 3.6 mm and 6.4 mm were the minimum unconstrained lengths required to survive the assumed lifetime of wind and snow cycles, respectively, confirming that typical industrial module constructions with 8–15 mm unconstrained lengths should survive conservatively. Notably, large magnitude, low-cycle snow loading was consistently the limiting factor requiring a longer unconstrained interconnect length. Finally, insights and workflows from this study inform module interconnection design limits for survival against mechanical fatigue in deployment environments.

14 SOLAR ENERGY↗

DOE OSTI2021

Probing function in 3D neuronal cultures: a survey of 3D multielectrode array advances

Recent advances in microphysiological systems (MPS) have made significant strides to include design features that reconstruct key elements found in the brain, and in parallel advanced technologies to detect the activity of electrogenic cells that form neural networks. In particular, three-dimensional multielectrode arrays (3D MEAs) are being developed with the increasing levels of spatial and temporal precision, difficult to achieve with current 2D MEAs, insertable MEA probes, and/or optical imaging of calcium dynamics. Thus, providing a means to monitor the flow of neural network activity within all three dimensions (X, Y, and Z) of the engineered tissue. In the last 6 years, 3D MEAs, using either “bottom-up” or “top-down” designs, have been developed to overcome the current technical challenges in monitoring the functionality of the in vitro systems. Herein, we will report on the design and application of novel 3D MEA prototypes for probing neural activity throughout the 3D neural tissue.

59 BASIC BIOLOGICAL SCIENCES↗

NASA NTRS1994

Vulnerability of manned spacecraft to crew loss from orbital debris penetration

Orbital debris growth threatens the survival of spacecraft systems from impact-induced failures. Whereas the probability of debris impact and spacecraft penetration may currently be calculated, another parameter of great interest to safety engineers is the probability that debris penetration will cause actual spacecraft or crew loss. Quantifying the likelihood of crew loss following a penetration allows spacecraft designers to identify those design features and crew operational protocols that offer the highest improvement in crew safety for available resources. Within this study, a manned spacecraft crew survivability (MSCSurv) computer model is developed that quantifies the conditional probability of losing one or more crew members, P(sub loss/pen), following the remote likelihood of an orbital debris penetration into an eight module space station. Contributions to P(sub loss/pen) are quantified from three significant penetration-induced hazards: pressure wall rupture (explosive decompression), fragment-induced injury, and 'slow' depressurization. Sensitivity analyses are performed using alternate assumptions for hazard-generating functions, crew vulnerability thresholds, and selected spacecraft design and crew operations parameters. These results are then used to recommend modifications to the spacecraft design and expected crew operations that quantitatively increase crew safety from orbital debris impacts.

Williamsen, J. E.↗

NASA NTRS2016

International Docking Standard (IDSS) Interface Definition Document (IDD) : Revision - E

This International Docking System Standard (IDSS) Interface Definition Document (IDD) is the result of a collaboration by the International Space Station membership to establish a standard docking interface to enable on-orbit crew rescue operations and joint collaborative endeavors utilizing different spacecraft. This IDSS IDD details the physical geometric mating interface and design loads requirements. The physical geometric interface requirements must be strictly followed to ensure physical spacecraft mating compatibility. This includes both defined components and areas that are void of components. The IDD also identifies common design parameters as identified in section 3.0, e.g., docking initial conditions and vehicle mass properties. This information represents a recommended set of design values enveloping a broad set of design reference missions and conditions, which if accommodated in the docking system design, increases the probability of successful docking between different spacecraft. This IDD does not address operational procedures or off-nominal situations, nor does it dictate implementation or design features behind the mating interface. It is the responsibility of the spacecraft developer to perform all hardware verification and validation, and to perform final docking analyses to ensure the needed docking performance and to develop the final certification loads for their application. While there are many other critical requirements needed in the development of a docking system such as fault tolerance, reliability, and environments (e.g. vibration, etc.), it is not the intent of the IDSS IDD to mandate all of these requirements; these requirements must be addressed as part of the specific developer's unique program, spacecraft and mission needs. This approach allows designers the flexibility to design and build docking mechanisms to their unique program needs and requirements. The purpose of the IDSS IDD is to provide basic common design parameters to allow developers to independently design compatible docking systems. The IDSS is intended for uses ranging from crewed to autonomous space vehicles, and from Low Earth Orbit (LEO) to deep-space exploration missions.The purpose of the IDSS IDD is to provide basic common design parameters to allow developers to independently design compatible docking systems. The IDSS is intended for uses ranging from crewed to autonomous space vehicles, and from Low Earth Orbit (LEO) to deep-space exploration missions. The purpose of the IDSS IDD is to provide basic common design parameters to allow developers to independently design compatible docking systems. The IDSS is intended for uses ranging from crewed to autonomous space vehicles, and from Low Earth Orbit (LEO) to deep-space exploration missions.

docking↗

NASA NTRS1987

Performance characteristics of moving belt radiators

A design features and performance capabilities evaluation is made for a novel, 'hybrid belt radiator' spacecraft heat rejection system, which retains the excellent heat transfer capacity of liquid belt radiators but does not require the direct exposure of a free liquid surface to the space environment. Attention is given to the preliminary results of a computer model analyzing the dynamic behavior of the flexible belt structure due to spacecraft accelerations, as well as to the results of system studies determining size constraints on the radiator. Over the 300-700 K range of operating temperatures, Space Shuttle-stowable radiators with 10-200 MW thermal capacities can be designed.

Mcfadden, D.↗

NASA NTRS1984

Tandem Fan Applications in Advanced STOVL Fighter Configurations

The series/parallel tandem fan engine is evaluated for application in advanced STOVL supersonic fighter aircraft. Options in engine cycle parameters and design of the front fan flow diverter are examined for their effects on engine weight, dimensions, and other factors in integration of the engine with the aircraft. Operation of the engine in high-bypass flow mode during cruise and loiter flight is considered as a means of minimizizng fuel consumption. Engine thrust augmentation by burning in the front fan exhaust is discussed. Achievement of very sort takeoff with vectored thrust in briefly reviewed for tandem fan engine configurations with vectorable front fan nozzles. Examples are given of two aircraft configuration planforms, a delta-canard, and a forward-swept wing, to illustrate the major features. design considerations, and potential performance of the tandem fan installation in each. Full realization of the advantages of tandem fan propulsion are found to depend on careful selection of the aircraft configuration, since integration requirements can strongly influence the engine performance.

Zola, Charlse L.↗

NASA NTRS1984

Tandem fan applications in advanced STOVL fighter configurations

The series/parallel tandem fan engine is evaluated for application in advanced STOVL supersonic fighter aircraft. Options in engine cycle parameters and design of the front fan flow diverter are examined for their effects on engine weight, dimensions, and other factors in integration of the engine with the aircraft. Operation of the engine in high-bypass flow mode during cruise and loiter flight is considered as a means of minimizing fuel consumption. Engine thrust augmentation by burning in the front fan exhaust is discussed. Achievement of very short takeoff with vectored thrust is briefly reviewed for tandem fan engine configurations with vectorable front fan nozzles. Examples are given of two aircraft configuration planforms, a delta-canard, and a forward-swept wing, to illustrate the major features, design considerations, and potential performance of the tandem fan installation in each. Full realization of the advantages of tandem fan propulsion are found to depend on careful selection of the aircraft configuration, since integration requirements can strongly influence the engine performance. Previously announced in STAR as N84-24579

Zola, C. L.↗

NASA NTRS1982

Future capabilities of the Delta launch vehicle

Design features and payload delivery capabilities of the Delta 3920 launch vehicle are presented. The 3920 is the 14th iteration in a series that was first launched in 1960 and has in recent years surpassed a 97 percent success rate. Capacity is 2800 lb to GEO. The second stage has a propellant capacity of 13,244 lb and a specific impulse of 319.2 sec and produces 9443 lb thrust. The first mission is the Landsat-D satellite and 27 more launches are scheduled through 1985, some associated with different upper stage configurations. The 3924 configuration delivers payload to GEO at a cost of $7500/lb. A fairing has been designed to provide back-up for Shuttle payloads, although the larger size reduces the GEO payload to 2715 lb. Performance enhancements may, however, be gained through pneumatic thrusters for the Castor IV engine separation, higher Castor IV thrust levels, increased booster thrust, and a cryogenic second stage.

Grimes, D. W.↗

NASA NTRS2006

Fuel Cells Utilizing Oxygen From Air at Low Pressures

A fuel cell stack has been developed to supply power for a high-altitude aircraft with a minimum of air handling. The fuel cell is capable of utilizing oxygen from ambient air at low pressure with no need for compression. For such an application, it is advantageous to take oxygen from the air (in contradistinction to carrying a supply of oxygen onboard), but it is a challenging problem to design a fuel-cell stack of reasonable weight that can generate sufficient power while operating at reduced pressures. The present fuel-cell design is a response to this challenge. The design features a novel bipolar plate structure in combination with a gas-diffusion structure based on a conductive metal core and a carbon gas-diffusion matrix. This combination makes it possible for the flow fields in the stack to have a large open fraction (ratio between open volume and total volume) to permit large volumes of air to flow through with exceptionally low backpressure. Operations at reduced pressure require a corresponding increase in the volume of air that must be handled to deliver the same number of moles of oxygen to the anodes. Moreover, the increase in the open fraction, relative to that of a comparable prior fuel-cell design, reduces the mass of the stack. The fuel cell has been demonstrated to operate at a power density as high as 105 W/cm2 at an air pressure as low as 2 psia (absolute pressure 14 kPa), which is the atmospheric pressure at an altitude of about 50,000 ft ( 15.2 km). The improvements in the design of this fuel cell could be incorporated into designs of other fuel cells to make them lighter in weight and effective at altitudes higher than those of prior designs. Potential commercial applications for these improvements include most applications now under consideration for fuel cells.

Cisar, Alan↗